Overall design method for simple model airplane in middle and primary schools
Through a systematic design method, including determining usage requirements, computing parameters and structural design, the problem of insufficient stability, endurance and safety of model aircraft is solved, and the overall performance of model aircraft is improved, which is suitable for teaching needs of primary and secondary schools.
Patent Information
- Application Number
- CN202411263735.3
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2024-09-10
- Publication Date
- 2025-02-18
AI Technical Summary
The lack of effective overall design methods in the prior art has led to insufficient stability, endurance and safety of model aircraft, which is difficult to meet the teaching needs of primary and secondary schools.
A general design method for simple model aircraft in primary and secondary schools is proposed, including determining the use requirements, calculating the first parameters (such as flight speed, takeoff weight, lift coefficient, etc.), and calculating the second parameters (such as wing parameters, secondary parameters, wing parameters, fuselage parameters and tail parameters), and finally determine the installation position of the wing on the fuselage and conduct structural design.
Through systematic design methods, the stability, endurance and safety of model aircraft are improved, making it more suitable for the teaching needs of primary and secondary schools.
Smart Images
Figure CN119475546A_ABST
Abstract
Description
Technical Field
[0001] The present application relates to the technical field of model aircraft design, and in particular to an overall design method for simple model aircraft for primary and secondary schools. Background Art
[0002] Currently, only real aircraft are known to have overall design methods. Model aircraft do not have a systematic design method. Their design is generally based on personal experience, and only some components have separate design methods.
[0003] The overall design of a real aircraft is relatively complex, and requires more than 20 major items and several minor items, such as determining requirements, evaluating plans, determining parameters, evaluating aircraft takeoff weight, and determining thrust. For example, determining requirements requires quality house analysis, fishbone diagram analysis, and R&D cost evaluation. When determining parameters, a large number of iterations are required to find the estimated value of aircraft takeoff weight, aircraft thrust and other parameters. Basic parameter calculations will also involve some calculus and complex data processing graphics. Model aircraft do not involve the same rigorous and detailed parameter requirements as real aircraft, and there is no complex demand analysis, complex theoretical calculations, and countless iterative calculations. Therefore, it is not realistic to substitute the design method of real aircraft into the design of model aircraft.
[0004] Although there are some toy model airplanes on the market that are designed based on experience, the lack of effective overall aircraft design methods has resulted in the lack of correlation between the various components of the model airplanes, which makes the model airplanes lack stability, endurance, and safety, making it difficult to meet the teaching needs of primary and secondary schools. Summary of the invention
[0005] In view of the above problems, the present application proposes an overall design method for simple model aircraft for primary and secondary schools, aiming to fill the gap in the overall design method of model aircraft and improve the stability, endurance and safety of model aircraft.
[0006] In order to achieve the above application objectives, this application provides the following technical solutions: A general design method for a simple model airplane for primary and secondary schools comprises the following steps: Determine the use requirements of model aircraft; Determine the first parameters of the model aircraft according to the use requirements, the first parameters including flight speed, take-off weight, lift coefficient, wing aspect ratio and wing tip root ratio; Calculating a second parameter of the model aircraft according to the first parameter, the second parameter comprising at least one of a wing parameter, an aileron parameter, a fuselage parameter, and a tail parameter; Calculate the center of gravity position of the model aircraft according to the first parameter and the second parameter; Determine the mounting position of the wing on the fuselage; Performing structural design according to the first parameter, the second parameter and the wing installation position; Make model airplanes.
[0007] As a preferred technical solution of the present application, the wing parameters include wing area, wing span, wing tip chord length and wing root chord length, among which, Calculate the wing area using the lift formula: (1) The wing span, wing tip chord length and wing root chord length are solved jointly using the following formula: (2), (3), AR = (4), TR = (5) In equations (1) to (5), L is the wing lift; ρ is the air density, ρ = 1.18 kg / m 3 ; V is the flight speed; S is the wing area; W is the takeoff weight; g is the acceleration due to gravity; Cl is the lift coefficient; C AW is the average geometric chord length; C r is the chord length of the wing tip; C t is the root chord length of the wing; b is the wing span; AR is the wing aspect ratio; TR is the wing tip root ratio.
[0008] As a preferred technical solution of the present application, the aileron parameters are selected according to the wing parameters, the aileron area is 5% to 7% of the wing area; and / or the aileron chord length is 20% to 25% of the wing chord length.
[0009] As the preferred technical solution of this application, the fuselage parameters include the fuselage length, and the fuselage length is calculated using the following formula: (6), where p is the wing-to-body ratio, and the value range of p is 1.03~1.75.
[0010] As a preferred technical solution of the present application, the fuselage parameters also include the middle width of the fuselage and the tail width of the fuselage, wherein the ratio of the fuselage length to the middle width of the fuselage ranges from 5.9 to 11.2; and / or the ratio of the fuselage length to the tail width of the fuselage ranges from 14 to 16.4.
[0011] As a preferred technical solution of the present application, the tail parameters include the horizontal tail area, and the horizontal tail area is calculated using the following formula: (7) In formula (7), C HT is the horizontal tail capacity, and its value range is 0.4~1.0; L Tis the tail arm, and its value range is 0.45I r ~0.55I r , S HT is the horizontal tail area.
[0012] As a preferred technical solution of the present application, the tail parameters also include the vertical tail area, and the vertical tail area is calculated using the following formula: (8) In formula (8), C VT is the vertical tail capacity, and its value range is 0.06~0.09; S VT is the vertical tail area.
[0013] As the preferred technical solution of this application, the center of gravity position of the model aircraft is calculated according to the following formula: (9) Among them, X is the center of gravity of the model aircraft, L 机翼 Calculated by formula (1), L 水平尾翼 =1 / 2×ρ×V 2 ×S HT ×CL.
[0014] As a preferred technical solution of the present application, the structural design is performed according to the first parameter, the second parameter and the wing installation position, including: Draw a top view of the model aircraft according to the first parameter, the second parameter and the wing installation position; It is judged whether the shape of the model aircraft meets the expected requirements according to the top view. If it does, the structure is modeled; if it does not, the second parameter is recalculated iteratively.
[0015] As a preferred technical solution of the present application, after manufacturing the model aircraft according to the modeling model, the design method further includes: According to the test flight results of the model aircraft, determine whether to iteratively recalculate the aircraft parameters.
[0016] Compared with the prior art, the design method of the present application has the following beneficial effects: First, the present application systematically proposes an overall design method for model aircraft, which solves the problems of complex parameters, huge amount of data calculation, high professional barriers, etc. caused by directly designing model aircraft according to the design method of real aircraft in the prior art. Secondly, compared with design based on experience, the overall design method proposed in the present application can effectively standardize, theorize and scientificize the various parameter indicators of model aircraft through the selection and calculation of multiple parameters, which is conducive to improving the stability, endurance and safety of model aircraft, and is conducive to the popularization and application of model aircraft in primary and secondary schools, providing convenience for primary and secondary school teaching. BRIEF DESCRIPTION OF THE DRAWINGS
[0017] Figure 1 A step diagram of a design method provided for some embodiments of the present application; Figure 2 Step diagrams of design methods provided in other embodiments of the present application. DETAILED DESCRIPTION
[0018] To make the purpose, technical solution and advantages of the embodiments of the present application clearer, the technical solution in the embodiments of the present application will be described clearly and completely in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present application, not all of them.
[0019] Therefore, the following detailed description of the embodiments of the present application is not intended to limit the scope of the present application for protection, but merely represents some embodiments of the present application. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present application.
[0020] It should be noted that, in the absence of conflict, the embodiments in this application and the features and technical solutions in the embodiments may be combined with each other.
[0021] It should be noted that similar reference numerals and letters denote similar items in the following drawings, and therefore, once an item is defined in one drawing, further definition and explanation thereof is not required in subsequent drawings.
[0022] In the description of the present application, it should be noted that the terms "upper", "lower", etc. indicate the orientation or positional relationship based on the orientation or positional relationship shown in the drawings, or the orientation or positional relationship in which the invention product is usually placed when in use, or the orientation or positional relationship commonly understood by those skilled in the art. Such terms are only for the convenience of describing the present application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore cannot be understood as a limitation on the present application. In addition, the terms "first", "second", etc. are only used to distinguish the description, and cannot be understood as indicating or implying relative importance.
[0023] Some preferred embodiments of the present application are described in detail below with reference to the accompanying drawings.
[0024] See also Figure 1 , the design method of this application includes the following steps: S1: Determine the use requirements of model aircraft; It should be understood that in step S1, the use requirements of the model aircraft should come from the designer's judgment after confirming the application scenario of the model aircraft. After the application scenario of the model aircraft is determined, the use requirements of the model aircraft can be estimated accordingly, such as flight speed, flight altitude, endurance level, etc.
[0025] S2: determining the first parameters of the model aircraft according to the use requirements, the first parameters including flight speed, take-off weight, lift coefficient, wing aspect ratio and wing tip root ratio; In the present application, the flight speed and take-off weight in the first parameter can be directly determined according to the use requirements, for example, according to the size of the flight field of the model aircraft, the safety of personnel in the flight field, and the shape of the model aircraft. The lift coefficient can be determined according to the wing airfoil and the wing installation angle. Under normal circumstances, model aircraft generally use flat airfoils, and the installation angle is generally below 10°. For example, the embodiment of the present application can take the installation angle as 6°, and the lift coefficient is determined to be 0.55. The wing aspect ratio and the wing tip root ratio can be selected according to the type of model aircraft, for example, can be determined by the following Table 1:
[0026] Table 1 S3: Calculating a second parameter of the model aircraft according to the first parameter, the second parameter including at least one of a wing parameter, an aileron parameter, a fuselage parameter, and a tail parameter; In the present application, wing parameters may also include wing area, wing span, wing tip chord length and wing root chord length, wherein the wing area can be solved by the lift calculation formula of the wing, as follows:
[0027] Where W is the takeoff weight, V is the flight speed, and both can be directly selected according to the use requirements of the model aircraft; g is the acceleration of gravity, usually g=9.8m / s 2 ; ρ is the air density. Taking room temperature of 25°C as an example, ρ=1.18Kg / m 3 ; Cl is the lift coefficient. In this application, Cl=0.55 can be selected; from this, the only unknown number S can be calculated, that is, the wing area S=0.618m 2 .
[0028] Furthermore, after the wing area S is obtained, the following formula can be used to solve the wing span, wing tip chord length and wing root chord length:
[0029]
[0030] AR =
[0031] TR =
[0032] In the formula, AR and TR are the wing aspect ratio and the wing tip root ratio, respectively. According to Table 1, the values of AR and TR can be selected according to the type of aircraft. For example, in the embodiment of the present application, AR=5 and TR=0.8 can be selected. Thus, the four unknowns C can be obtained by combining the four equations: r , C t , C AW , b, that is, the chord length of the wing tip C r =0.164m, wing root chord length C t =0.2m, average geometric chord length C AW =0.182m, b=0.91m.
[0033] In the present application, aileron parameters may include aileron area and aileron chord length. Usually, there are no strict requirements for aileron parameters, and they can be selected based on experience. For example, in general, the aileron area can be 5% to 7% of the wing area, and the aileron chord length can be 20% to 25% of the wing chord length.
[0034] In this application, the fuselage parameters at least include the fuselage length, which can be calculated using the following formula:
[0035] Where p is the wing-to-body ratio. The value of p can be selected by referring to Table 2 below.
[0036] Table 2 It can be clearly seen from Table 2 that the wing-to-body ratio p can also be selected according to the type of model aircraft. For example, in the embodiment of the present application, p can be selected as 1.38, and the fuselage length I can be calculated from this. r = 0.66m.
[0037] Furthermore, the fuselage parameters may also include the width of the middle fuselage and the width of the tail fuselage. When the fuselage length is known, the ratio of the aircraft length to the width may be empirically selected according to Table 3 below, thereby calculating the width of the middle fuselage and the width of the tail fuselage. For example, in the embodiment of the present application, the ratio of the fuselage length to the width of the middle fuselage may be selected as 6, and the ratio of the fuselage length to the width of the tail fuselage may be selected as 15, thereby calculating that the width of the middle fuselage is 0.11m and the width of the tail fuselage is 0.04m.
[0038]
[0039] Table 3 In this application, the aircraft tail parameters include horizontal tail parameters and vertical tail parameters, and the most important of the horizontal tail parameters is the horizontal tail area, which can be calculated using the following formula:
[0040] In the formula, C HT is the horizontal tail capacity, which can be selected empirically according to Table 4 below; L T is the tail lever, which can be based on the percentage of the fuselage length. For an aircraft with an engine mounted on the wing, the tail lever is about 50%-55% of the fuselage length. For an aircraft with an engine mounted at the rear, the tail lever is about 45%-50% of the fuselage length. HT The value is 0.3, and L T =0.55×I r , from which, substituting into the calculation, we can get the horizontal tail area S HT =0.025m 2 .
[0041]
[0042] Table 4 Further, the horizontal tail area S is calculated HT After that, the span b of the horizontal tail can be further calculated H , tip string length C rH and the root chord length C tH The calculation formulas of these three are the same as the calculation formulas of wing span, wing tip chord length and wing root chord length, which will not be repeated here. The difference is that the empirical values of aspect ratio and tip-to-root ratio of the horizontal tail are different from those of the wing. In this application, the empirical values of aspect ratio and tip-to-root ratio of the horizontal tail can be selected according to the following Table 5:
[0043] Table 5 In the embodiment of the present application, the aspect ratio of the horizontal tail can be 3, and the tip-to-root ratio of the horizontal tail can be 0.4. Then, b can be calculated by the simultaneous formula: H =0.27m, C rH =0.06m, C tH =0.12m.
[0044] Similar to the horizontal tail parameters, the vertical tail parameters also include the vertical tail area, vertical tail span, vertical tail tip chord length and vertical tail root chord length. First, the vertical tail area is calculated according to the following formula:
[0045] In the formula, C VT is the vertical tail capacity, which can be selected empirically according to Table 4 above; L T is the tail lever, which can be based on the percentage of the fuselage length. For an aircraft with an engine mounted on the wing, the tail lever is about 50%-55% of the fuselage length. For an aircraft with an engine mounted at the rear, the tail lever is about 45%-50% of the fuselage length. VT The value is 0.08, and L T =0.55×I r , thus, substituting into the calculation, we can get the vertical tail area S VT =0.0337m 2 .
[0046] Similarly, the span b of the vertical tail V , tip string length C rV and the root chord length C tV The calculation formulas of these three are also the same as the simultaneous calculation formulas of wing span, wing tip chord length and wing root chord length, which will not be repeated here. The empirical values of the aspect ratio and tip-to-root ratio of the vertical tail can be selected according to Table 5 above. For example, in the embodiment of the present application, the aspect ratio of the vertical tail can be 2, and the tip-to-root ratio of the vertical tail can be 0.4. From this, it can be calculated by the simultaneous formula that b V =0.36m, C rV =0.11m, C tV =0.25m. So far, the second parameters of the model aircraft have all been calculated.
[0047] S4: Calculating the center of gravity position of the model aircraft according to the first parameter and the second parameter; It is particularly noted that, under normal circumstances, the relative position of the fuselage and the tail is generally easy to determine, but the relative position relationship between the wing and the fuselage is difficult to determine. Therefore, the present application proposes to determine the position of the wing on the fuselage by finding the center of gravity of the aircraft. The center of gravity of a model aircraft is not determined by a large number of formulas and wind tunnel tests like a real aircraft, but can be determined by some experience, that is, the center of gravity of a general model aircraft is one-third of the distance from the nose and one-third of the leading edge of the wing. By setting an aircraft balance equation based on this relationship, the position of the wing can be determined. When the aircraft as a whole tilts a few degrees, the torque generated by the wing is equal to the torque generated by the horizontal tail. Specifically, the center of gravity of the aircraft is Wing position, the lift point generated by the wing and horizontal tail is Position, the overall angle of attack of the aircraft is balanced, and the horizontal tail arm is set to X, the equilibrium equation can be obtained: Where, L 机翼 It can be obtained from the wing lift formula, L水平尾翼 The calculation method is similar to the wing lift formula, L 水平尾翼 =1 / 2×ρ×V 2 ×S HT ×CL, after substituting all the data into the calculation, we can get X=0.353m, from which the center of gravity of the model aircraft can be calculated.
[0048] S5: Determine the installation position of the wing on the fuselage.
[0049] In step S5, it can be understood that after determining the center of gravity position of the model aircraft, the installation position of the wing can be easily determined. Generally, the position of the center of gravity of the empty aircraft weight is about 30% of the front area of the wing.
[0050] S6: Perform structural design according to the first parameter, the second parameter and the wing installation position.
[0051] See also Figure 2 In step S6, performing structural design according to the first parameter, the second parameter and the wing installation position may also include: S61: Draw a top view of the model aircraft according to the first parameter, the second parameter and the wing installation position; S62: judging whether the shape of the model aircraft meets the expected requirements according to the top view, and if so, performing structural modeling; if not, re-iteratively calculating the second parameter.
[0052] It is understandable that, in step S61, the top view of the model aircraft can be drawn by using some basic design software, such as the commonly used CATIA design software.
[0053] In step S62, firstly, it is necessary to roughly observe whether the structure of the model aircraft is reasonable and whether it meets the expected assumptions according to the top view of the model aircraft. If it meets the expectations, subsequent structural modeling can be carried out. The structural modeling can also be carried out using CATIA design software. When modeling, it is necessary to consider the matching design between various parts of the aircraft, such as slot design, installation method design, radian design, etc. If it is found that the structure has obvious defects or does not meet the expected assumptions after observing the top view, the second parameter should be re-iterated and calculated, and the focus should be on whether the selection of some parameter experience values is reasonable, and then step S6 should be re-executed according to the new calculation results.
[0054] S7: Make model airplanes.
[0055] In step S7, the model airplane can be made by hand or by machine, or some parts can be made by hand and some parts can be machined, and this application does not limit this. The material of the model airplane can be hard cardboard material, plastic material, or part of the metal material, and this application does not limit this.
[0056] Furthermore, in some embodiments, the design method of the present application may further include step S8: S8: Determine whether to iteratively recalculate aircraft parameters based on the test flight results of the model aircraft.
[0057] By testing the model aircraft, the flight speed, endurance, flight stability, etc. of the model aircraft can be more intuitively and accurately determined, and whether it is necessary to re-iterate the calculation according to the test flight performance. If it is not necessary to iterate the calculation, the design of the model aircraft is completed; if it is necessary to improve certain performances of the model aircraft, the specific parameters that affect the performance of the model aircraft must first be determined, and then these parameters must be iteratively optimized and calculated, so as to re-execute steps S6 to S8.
[0058] The design method of the present application has at least the following beneficial effects: First, the present application systematically proposes an overall design method for model aircraft, which solves the problems of complex parameters, huge amount of data calculation, high professional barriers, etc. caused by directly designing model aircraft according to the design method of real aircraft in the prior art. Secondly, compared with design based on experience, the overall design method proposed in the present application can effectively standardize, theorize and scientificize the various parameter indicators of model aircraft through the selection and calculation of multiple parameters, which is conducive to improving the stability, endurance and safety of model aircraft, and is conducive to the popularization and application of model aircraft in primary and secondary schools, providing convenience for primary and secondary school teaching.
[0059] The present application also proposes a model aircraft, which is designed using the design method of any of the above embodiments.
[0060] The above embodiments are only used to illustrate the present application but not to limit the technical solutions described in the present application. Although the present application has been described in detail with reference to the above embodiments, the present application is not limited to the above specific implementation methods. Therefore, any modification or equivalent replacement of the present application; and all technical solutions and improvements thereof that do not depart from the spirit and scope of the invention are included in the scope of the claims of the present application.
Claims
1. A general design method for a simple model airplane for primary and secondary schools, characterized in that: The following steps are involved: Determine the use requirements of model aircraft; Determining first parameters of the model aircraft according to usage requirements, wherein the first parameters include flight speed, take-off weight, lift coefficient, wing aspect ratio, and wing tip root ratio; Calculating a second parameter of the model aircraft according to the first parameter, the second parameter comprising at least one of a wing parameter, an aileron parameter, a fuselage parameter, and a tail parameter; Calculating the center of gravity of the model aircraft according to the first parameter and the second parameter; Determine the mounting position of the wing on the fuselage; Performing structural design according to the first parameter, the second parameter and the wing installation position; Make model airplanes.
2. The design method according to claim 1, characterized in that: The wing parameters include wing area, wing span, wing tip chord length and wing root chord length, wherein: Calculate the wing area using the lift formula: （1） The wing span, the wing tip chord and the wing root chord are jointly solved using the following formula: （2）， （3）， AR = (4), TR = (5) In equations (1) to (5), L is the wing lift; ρ is the air density, ρ = 1.18 kg / m 3 ; V is the flight speed; S is the wing area; W is the takeoff weight; g is the acceleration of gravity; Cl is the lift coefficient; C AW is the average geometric chord length; C r is the chord length of the wing tip; C t is the root chord length of the wing; b is the wing span; AR is the wing aspect ratio; TR is the wing tip root ratio.
3. The design method according to claim 2, characterized in that: The aileron parameters are selected according to the wing parameters, wherein, Aileron area is 5% to 7% of the wing area; and / or The aileron chord length is 20%~25% of the wing chord length.
4. The design method according to claim 2, characterized in that: The fuselage parameters include fuselage length, and the fuselage length is calculated using the following formula: （6） Among them, p is the wing-to-body ratio, and the value range of p is 1.03~1.
75.
5. The design method according to claim 4, characterized in that: The fuselage parameters also include the fuselage middle width and the fuselage tail width, wherein: The ratio of the fuselage length to the fuselage middle width is in the range of 5.9 to 11.2; and / or The ratio of the fuselage length to the fuselage tail width ranges from 14 to 16.
4.
6. The design method according to claim 4, characterized in that: The tail parameters include the horizontal tail area, and the horizontal tail area is calculated using the following formula: （7） In formula (7), C HT is the horizontal tail capacity, and its value range is 0.4~1.0; L T is the tail arm, and its value range is 0.45I r ~0.55I r , S HT is the horizontal tail area.
7. The design method according to claim 6, characterized in that: The tail parameters also include the vertical tail area, which is calculated using the following formula: （8） In formula (8), C VT is the vertical tail capacity, and its value range is 0.06~0.09; S VT is the vertical tail area.
8. The design method according to claim 6, characterized in that: The center of gravity of the model aircraft is calculated according to the following formula: （9） Among them, X is the center of gravity of the model aircraft, L 机翼 Calculated by formula (1), L 水平尾翼 =1 / 2×ρ×V 2 ×S HT ×CL.
9. The design method according to claim 1, characterized in that: The performing structural design according to the first parameter, the second parameter and the wing installation position comprises: Draw a top view of the model aircraft according to the first parameter, the second parameter and the wing installation position; It is judged whether the shape of the model aircraft meets the expected requirements according to the top view. If it does, the structure is modeled; if it does not, the second parameter is recalculated iteratively.
10. The design method according to claim 1, characterized in that: After manufacturing the model aircraft according to the modeling model, the design method further comprises: According to the test flight results of the model aircraft, determine whether to iteratively recalculate the aircraft parameters.